BACKGROUND OF THE INVENTION
[0001] This invention relates generally to composite materials, and more specifically to
composite panels having a honeycomb core that are manufactured using a single-step
curing process.
[0002] Honeycomb core sandwich panels include composite laminate skins co-cured with adhesives
to a honeycomb core are known. Such panels are utilized, for example, in the aerospace
industry. Such honeycomb core sandwich panels are used at least partially due to their
high stiffness-to-weight (i.e., specific stiffness) and strength-to-weight (i.e.,
specific strength) ratios. Current sandwich panels are processed, for example, using
multiple cure cycles, or are co-cured in one step at low cure pressures in order to
prevent deformation of the honeycomb core.
[0003] When multiple cure cycles are used, either one or both of the composite skins are
precured, then later bonded or cobonded onto the honeycomb core and/or the other skin.
The multiple and complex steps increase production expenses and also increase the
risk of defective panels due to, for example, bondline discrepancies, since the precured
skin must match very closely with the contour-machined honeycomb core surface.
[0004] A co-cure process refers to more than one skin simultaneously cured and bonded onto
the machined honeycomb core pieces. A skin may include one ply of a composite material
or a plurality of plies of a composite material. Numerous quality issues typically
occur with a co-cure process such as, for example, distortion of the composite plies
and/or dispersion of the composite resin into cells of the honeycomb core, as well
as distortion of cells of the honeycomb core from their pre-processed core cell shape.
Other quality issues due to known co-cure processes may also include formation of
porous composite skins caused by curing at a less than optimum pressure, and also
due to distortion of the composite skins into the cells of the core.
[0005] Furthermore, quality inspections of composite panels formed using known co-cure processes
are complicated because it may be difficult to accurately assess the quality of a
composite panel formed using a known co-cure process using typical non-destructive
inspection techniques, such as ultrasonic inspection methods.
[0006] There are no current composite/honeycomb sandwich design and process schemes in which
a one-step co-cure process produces a high strength structural part, with substantially
porosity-free composite skins, minimal resin leakage into the honeycomb core, minimal
honeycomb core distortion, minimal composite ply distortion into the honeycomb cells,
and having composite skins that may be readily inspected by known ultrasonic inspection
techniques to verify the co-cured skins are substantially porosity-free.
[0007] U.S. Patent Publication 2005/0161154 discloses methods of stabilizing and/or sealing core material for manufacturing molded
composite structures by applying a layer of roll-coated adhesive on the core, applying
a layer of stabilizing material on the roll-coated adhesive layer, and applying a
layer of thermoplastic barrier film on the stabilizing material.
[0008] EP 1 897 680 A1 discloses a method for producing a sandwich component having a honeycomb core, wherein
a barrier layer is provided between the core and the composite cover layers.
[0009] EP 0 786 330 A2 discloses low-weight and water-resistant honeycomb sandwich panels which are made
by a resin transfer molding process. A moisture barrier film is placed between the
core and a fibrous reinforcement material.
BRIEF DESCRIPTION OF THE INVENTION
[0010] In one embodiment, a method of manufacturing a composite panel is provided. The method
includes pre-stabilizing a honeycomb core by surrounding the honeycomb core with pre-stabilizing
material and curing the pre-stabilizing material; positioning a laminate skin formed
from one sheet of a composite material and folded along an edge thereof to surround
said pre-stabilized honeycomb core; bonding said at least one laminate skin to opposite
sides of said pre-stabilizing material surrounding said pre-stabilized honeycomb core;
and curing said at least one laminate skin to opposite sides of said pre-stabilizing
material substantially simultaneously.
[0011] In another embodiment, a composite panel including a pre-stabilized honeycomb core
and at least one laminate skin is provided. A composite panel comprising: a pre-stabilized
honeycomb core surrounded by a cured pre-stabilizing material; and a laminate skin
formed from one sheet of a composite material and folded along an edge thereof, said
one laminate skin positioned to surround said pre-stabilized honeycomb core, characterized
in that said laminate skin is bonded to opposite sides of said pre-stabilizing material
surrounding said pre-stabilized honeycomb core and cured substantially simultaneously.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012]
Figure 1 is an illustration of an exemplary composite panel
Figure 2 is a cross-sectional view of the exemplary composite panel of Figure 1 taken
generally along line 2-2.
Figure 3 is a flow chart illustrating an exemplary method of manufacturing the composite
panel of Figures 1 and 2.
DETAILED DESCRIPTION OF THE INVENTION
[0013] Figure 1 is an illustration of an exemplary composite panel 10. Exemplary composite
panel 10 is a sandwich structure that includes two core portions (not shown in Figure
1), positioned between, and surrounded by, two layers of composite laminate skin,
a first layer 16 and a second layer 18. In the exemplary embodiment, first laminate
skin layer 16 and second laminate skin layer 18 may be two separate sheets of composite
material. In an alternative embodiment, first laminate skin layer 16 and second laminate
skin layer 18 are formed from one sheet of composite material, and folded along an
edge, for example, but not limited to an edge 22 or an edge 24, such that first layer
16 is adjacent to one side of the core portions (not shown in Figure 1) and second
layer 18 is adjacent to an opposite side of the core portions (not shown in Figure
1).
[0014] Figure 2 is a cross-sectional view of exemplary composite panel 10 taken generally
along line 2-2. As described above, composite panel 10 includes at least one core
section 40, a pre-stabilizing composite material 42, and first and second composite
laminate layers 16 and 18. Examples of composite materials may include, but are not
limited to, fiberglass and carbon fiber in a resin matrix.
[0015] Core section 40 may be, for example, a honeycomb structure. Core section 40 may be
formed from, for example, but not limited to, paper, synthetic paper (e.g., NOMEX®
brand fiber, manufactured by E.I du Pont de Nemours and Company), metal, composite,
fiberglass, or the like, or any material that enables composite panel 10 to function
as described herein. For example, core section 40 may be a chamfered-edge honeycomb
core having a density of 3-8 pounds per cubic foot. In the exemplary embodiment, core
section 40 is pre-stabilized with pre-stabilizing composite material 42, resulting
in a pre-stabilized core 44. More specifically, core section 40 may be pre-stabilized
by coupling composite material 42, such as, but not limited to, a ply of preimpregnated
fiberglass, to core section 40. In the exemplary embodiment, composite material 42
is coupled to core section 40 using an adhesive 50, such as, but not limited to, a
film adhesive, and curing by heating. In an alternative embodiment, composite material
42 is coupled directly to core section 40.
[0016] First laminate skin layer 16 and second laminate skin layer 18 may include, but are
not limited to including, composite laminate materials such as carbon fiber or fiberglass.
First laminate skin layer 16 and second laminate skin layer 18 may be coupled to pre-stabilized
core section 40 by respective adhesive layers 52 and 54.
[0017] Figure 3 is a flow chart 60 illustrating an exemplary method of manufacturing the
composite panel 10 illustrated in Figures 1 and 2. The method includes pre-stabilizing
64 a core. In the exemplary embodiment, pre-stabilizing 64 includes applying a pre-stabilizing
composite material to the core and curing the pre-stabilizing composite material in
an oven cure cycle under a pressure. The pressure may be applied using, for example,
but not limited to, a vacuum bag or an autoclave. Pre-stabilizing 64 increases the
strength of the core section 40 without substantially increasing its weight. In the
exemplary embodiment, the core is machined to a desired shape and size prior to pre-stabilizing
64. In an alternative embodiment, the core is formed into a desired shape during pre-stabilizing
64. The applied heat facilitates changing the shape of the core and the pre-stabilizing
composite material maintains the desired shape of the core.
[0018] The method further includes positioning 66 the pre-stabilized core adjacent to at
least a first skin layer and a second skin layer. As described above in regard to
Figure 1, laminate skin layer 16 and laminate skin layer 18 may be formed from two
separate sheets of composite material. In an alternative embodiment, laminate skin
layer 16 and laminate skin layer 18 are formed from one sheet of composite material,
and folded along an edge, for example, but not limited to edge 22 or edge 24, such
that skin layer 16 is adjacent to one side of the cores and skin layer 18 is adjacent
to an opposite side of the cores.
[0019] Additionally, the method still further includes curing 68 the first skin layer 16
and the second skin layer 18 substantially simultaneously. In the exemplary embodiment,
curing 68 is a high pressure (e.g., 3.44-8.27 bar (50-120 psi)), one-step co-cure
process which cures both first skin layer 16 and second skin layer 18 while simultaneously
bonding first skin layer 16 and second skin layer 18 to a core, for example, core
section 40.
[0020] In the exemplary embodiment, curing 68 enables production of cured skins including
high-strength composite materials. Curing process 68, and more specifically the high
pressure curing that is facilitated by pre-stabilizing the core, results in cured
skins having a substantially low porosity, for example, a porosity of less than 2%,
which is one indication of high structural performance and of a high level of quality.
Curing 68 also enables production of cured skins that are inspectable by known non-destructive
inspection techniques, including, but not limited to, ultrasonic inspection techniques.
In the exemplary embodiment, the inspection techniques have a resolution sufficient
to quantify the porosity of the cured skins.
[0021] Furthermore, curing 68 and pre-stabilizing 64 facilitate prevention of core distortion
and prevention of resin from the skins intruding into the core during the cure. Pre-stabilizing
64 enables a composite panel to include a lightweight core while also being cured
at a high pressure, without the core being distorted. Curing 68 at a high pressure
produces skins having a low porosity.
[0022] The methods and apparatus described herein facilitate a one-step co-cure process
which results in a high strength structural part, that includes substantially porosity-free
composite skins, substantially no resin leakage into the honeycomb core, substantially
no honeycomb core distortion, substantially no composite ply distortion into the honeycomb
cells, while the sandwich skins are readily inspectable by known ultrasonic inspection
techniques to verify the co-cured skins are substantially porosity-free.
[0023] The methods and apparatus described herein facilitate cost-effective manufacturing
of high strength toughened carbon/epoxy composites. Specifically, the methods and
apparatus described herein facilitate increasing structural performance of composite
panels, minimizing weight, improving part quality inspectability, while minimizing
manufacturing cost (in part due to the one-step co-cure process described above).
[0024] Exemplary embodiments of composite panels and methods of manufacturing composite
panels are described above in detail. Neither the composite panels nor the methods
of manufacturing the composite panels are limited to the specific embodiments described
herein. Each component can also be used in combination with other components. More
specifically, although the methods and apparatus herein are described with respect
to aircraft components, it should be appreciated that the methods and apparatus can
also be applied to a wide variety of components used within other structures, including,
but not limited to, spacecraft, watercraft, and automobiles.
1. A composite panel (10) comprising:
a pre-stabilized honeycomb core (44) surrounded by a cured pre-stabilizing material;
and
a laminate skin (16,18) formed from one sheet of a composite material and folded along
an edge (22, 24) thereof, said one laminate skin (16, 18) positioned to surround said
pre-stabilized honeycomb core, characterized in that
said laminate skin (16, 18) is bonded to opposite sides of said pre-stabilizing material
(42) surrounding said pre-stabilized honeycomb core (44) and cured substantially simultaneously.
2. A composite panel (10) according to Claim 1 wherein said honeycomb core (44) is pre-stabilized
by applying a composite material (42) to said honeycomb core and curing said composite
material.
3. A composite panel (10) according to Claim 1 wherein said pre-stabilizing material
is coupled to said honeycomb core (44) using an adhesive (50).
4. A composite panel (10) according to Claim 1 wherein said at least one laminate skin
(16,18) is coupled to said pre-stabilized honeycomb core (44) using an adhesive (50).
5. A composite panel (10) according to Claim 1 wherein said laminate skin (16,18) comprises
at least one composite material (42).
6. A method of manufacturing a composite panel (10) comprising:
pre-stabilizing a honeycomb core (40) by surrounding the honeycomb core (40) with
pre-stabilizing material (42) and curing the pre-stabilizing material;
positioning a laminate skin (16,18) formed from one sheet of a composite material
and folded along an edge (22, 24) thereof to surround said pre-stabilized honeycomb
core (44);
bonding said at least one laminate skin (16,18) to opposite sides of said pre-stabilizing
material (42) surrounding said pre-stabilized honeycomb core (44); and curing said
at least one laminate skin (16,18) to opposite sides of said pre-stabilizing material
substantially simultaneously.
7. A method of manufacturing a composite panel (10) according to Claim 6 wherein said
laminate_skin (16, 18) of said composite panel is able to be inspected with at least
one non-destructive inspection method to quantify porosity.
8. A method of manufacturing a composite panel (10) according to Claim 7 wherein said
at least one non-destructive inspection method comprises ultrasonic testing.
1. Verbundplatte (10), die aufweist:
einen vorstabilisierten Wabenkern (44), der von einem ausgehärteten vorstabilisierten
Material umgeben ist; und
eine Laminathaut (16, 18), die aus einer Schicht eines Verbundmaterials gebildet ist
und die entlang eines Rands (22, 24) davon gefaltet ist, wobei die Laminathaut (16,
18) positioniert ist, den vorstabilisierten Wabenkern zu umgeben,
dadurch gekennzeichnet, dass
die Laminathaut (16, 18) an gegenüberliegende Seiten des vorstabilisierten Materials
(42) gebunden ist, das den vorstabilisierten Wabenkern (44) umgibt, und im Wesentlichen
gleichzeitig ausgehärtet wird.
2. Verbundplatte (10) nach Anspruch 1, wobei der Wabenkern (44) vorstabilisiert wird,
indem ein Verbundmaterial (42) auf den Wabenkern angewendet wird und indem das Verbundmaterial
ausgehärtet wird.
3. Verbundplatte (10) nach Anspruch 1, wobei das vorstabilisierte Material unter Verwendung
eines Klebstoffs (50) an den Wabenkern (44) gekoppelt ist.
4. Verbundplatte (10) nach Anspruch 1, wobei die zumindest eine Laminathaut (16, 18)
unter Verwendung eines Klebstoffs (50) an den vorstabilisierten Wabenkern (44) gekoppelt
ist.
5. Verbundplatte (10) nach Anspruch 1, wobei die Laminathaut (16, 18) zumindest ein Verbundmaterial
(42) aufweist.
6. Verfahren zum Herstellen einer Verbundplatte (10), das aufweist:
Vorstabilisieren eines Wabenkerns (40), indem der Wabenkern (40) mit einem vorstabilisierten
Material (42) umgeben wird und indem das vorstabilisierte Material ausgehärtet wird;
Positionieren einer Laminathaut (16, 18), die aus einer Schicht eines Verbundmaterials
gebildet ist und die entlang eines Rands (22, 24) davon gefaltet ist, um den vorstabilisierten
Wabenkern (44) zu umgeben;
Binden der zumindest einen Laminathaut (16, 18) an gegenüberliegende Seiten des vorstabilisierten
Materials (42), das den vorstabilisierten Wabenkern (44) umgibt; und
im Wesentlichen gleichzeitiges Aushärten der zumindest einen Laminathaut (16, 18)
an gegenüberliegenden Seiten des vorstabilisierten Materials.
7. Verfahren zum Herstellen einer Verbundplatte (10) nach Anspruch 6, wobei die Laminathaut
(16, 18) der Verbundplatte mit zumindest einem nicht zerstörerischen Inspektionsverfahren
inspizierbar ist, um eine Porosität zu quantifizieren.
8. Verfahren zum Herstellen einer Verbundplatte (10) nach Anspruch 7, wobei das zumindest
eine nicht zerstörerische Verfahren eine Ultraschallüberprüfung aufweist.
1. Panneau composite (10) comprenant :
une âme en nid d'abeilles (44) pré-stabilisée entourée par un matériau pré-stabilisant
durci ; et
une peau stratifiée (16, 18) formée à partir d'une feuille d'un matériau composite
et pliée le long d'un bord (22, 24) de celle-ci, ladite peau stratifiée (16, 18) étant
positionnée de façon à entourer ladite âme en nid d'abeilles pré-stabilisée,
caractérisé en ce que
ladite peau stratifiée (16, 18) est assemblée à des côtés opposés dudit matériau pré-stabilisant
(42) entourant ladite âme en nid d'abeilles (44) pré-stabilisée et durcie de façon
essentiellement simultanée.
2. Panneau composite (10) selon la revendication 1, dans lequel ladite âme en nid d'abeilles
(44) est pré-stabilisée par application d'un matériau composite (42) à ladite âme
en nid d'abeilles et durcissement dudit matériau composite.
3. Panneau composite (10) selon la revendication 1, dans lequel ledit matériau pré-stabilisant
est accouplé à ladite âme en nid d'abeilles (44) au moyen d'un adhésif (50).
4. Panneau composite (10) selon la revendication 1, dans lequel ladite ou lesdites peaux
stratifiées (16, 18) sont accouplées à ladite âme en nid d'abeilles (44) pré-stabilisée
au moyen d'un adhésif (50).
5. Panneau composite (10) selon la revendication 1, dans lequel ladite peau stratifiée
(16, 18) comprend au moins un matériau composite (42).
6. Procédé de fabrication d'un panneau composite (10) comprenant :
pré-stabiliser une âme en nid d'abeilles (40) en entourant l'âme en nid d'abeilles
(40) d'un matériau pré-stabilisant (42) et en durcissant le matériau pré-stabilisant;
positionner une peau stratifiée (16, 18) formée à partir d'une feuille d'un matériau
composite et pliée le long d'un bord (22, 24) de celle-ci de façon à entourer ladite
âme en nid d'abeilles (44) pré-stabilisée ;
assembler ladite ou lesdites peaux stratifiées (16, 18) à des côtés opposés dudit
matériau pré-stabilisant (42) entourant ladite âme en nid d'abeilles (44) pré-stabilisée
; et
durcir ladite ou lesdites peaux stratifiées (16, 18) sur des côtés opposés dudit matériau
pré-stabilisant de façon essentiellement simultanée.
7. Procédé de fabrication d'un panneau composite (10) selon la revendication 6, dans
lequel ladite peau stratifiée (16, 18) dudit panneau composite peut être inspectée
à l'aide d'au moins un procédé d'inspection non destructif afin de quantifier la porosité.
8. Procédé de fabrication d'un panneau composite (10) selon la revendication 7, dans
lequel ledit ou lesdits procédés d'inspection non destructifs comprennent l'auscultation
dynamique.